1. Field of the Invention
The present invention relates generally to fluid reaction surfaces and more specifically to turbine airfoils with film cooling.
2. Description of the Related Art Including Information Disclosed Under 37 CFR 1.97 and 1.98
A gas turbine engine includes a turbine section in which a hot gas flow passes through the rotor blades and stationary vanes or nozzles to extract energy from the flow. The efficiency of the gas turbine engine can be increased by providing a higher flow temperature into the turbine. However, the temperature is limited to the material capabilities of the parts exposed to the hot gas flow.
One method of allowing for higher turbine temperatures is to provide cooling of the first and second stages of the turbine. Complex internal cooling passages have been disclosed to provide high cooling capabilities while using lower flow volumes of cooling air. Since the cooling air used in the turbine airfoils is typically bleed off air from the compressor, using less bleed off air for cooling also increases the efficiency of the engine.
A Prior Art airfoil leading edge is cooled with backside impingement in conjunction with a showerhead film cooling and is shown in FIG. 1. The airfoil 10 includes a cooling air supply cavity 22, a metering hole 23 connecting the supply cavity 22 to an impingement cavity 24, and a plurality of impingement holes 27 to discharge cooling air to the leading edge surface of the airfoil 10. a mid-chord three-pass forward flowing serpentine cooling circuit includes a first leg channel 40, a second leg channel 36, and a third leg channel 34. Film cooling holes 35 deliver cooing air to the suction side and film cooling holes 37 deliver cooling air to the pressure side from the third leg channel 34. Turbulators 50 are positioned along the walls of the channels and cavity to promote heat transfer to the cooling air. A trailing edge cooling slots 42 and cooling air exit holes 44 discharge cooling air from the first leg channel 40 to the trailing edge of the airfoil.
The showerhead film rows are fed cooling air from a common cooling supply cavity 22 and discharged at various gas side pressures. The pressure at each of the gas side locations can vary substantially as the hot gas flow accelerates around the nose of the leading edge. The minimum pressure ratio across the showerhead holes 27 is typically set by back-flow margin requirements, and the pressure ratio (and flow) across all of the other film cooling rows becomes substantially a function of the gas-side pressure. Backflow occurs when the pressure of the hot gas flow outside the leading edge is higher than the cooling air pressure inside the cooling supply cavity 24, resulting in the hot gas flowing into the inside of the airfoil. As a result of this cooling design, the cooling flow distribution and pressure ratio across the showerhead film holes 27 for the pressure side and suction side film row is predetermined by the supply pressure.
U.S. Pat. No. 7,011,502 B2 issued to Lee et al on Mar. 14, 2006 entitled THERMAL SHIELD TURBINE AIRFOIL. Discloses an airfoil with a longitudinal first inlet channel (56 in this patent) connected by a row of impingement holes (48 in this patent) to a longitudinal channel (42 in this patent), being connected by a plurality of film cooling holes (50 in this patent) to the leading edge surface of the airfoil. In the Lee et al patent, the longitudinal channel is also connected to bridge channels on the pressure side and suction sides of the longitudinal channel. Only one multi-impingement channel is used in the Lee et al invention to supply film cooling holes as opposed to the three separate multi-impingement channels of the present invention.
U.S. Pat. No. 4,859,147 issued to Hall et al on Aug. 22, 1989 entitled COOLED GAS TURBINE BLADE discloses an airfoil with a showerhead arrangement having a cooling air supply cavity (24 in this patent), an impingement cavity (28 in this patent) connected to the cooling supply cavity by three slots (26 in this patent), and three film cooling holes (30 in this patent) connected to the impingement cavity.
U.S. Pat. No. 6,379,118 B2 issued to Lutum et al on Apr. 30, 2002 entitled COOLED BLADE FOR A GAS TURBINE discloses a similar showerhead arrangement to that above of the Hall et al patent. A cooling air supply cavity (50 in this patent) is connected to an impingement cavity (47 in this patent) through two impingement cooling holes (49 in this patent), and three film cooling holes (48 in this patent) are connected to the impingement cavity. Both of the Hall et al and Lutum et al patents lack the first impingement cavity and the second impingement cavity in series, and the multi-impingement cavities of the present invention.
It is an object of the present invention to alleviate the problem associated with the turbine airfoil leading edge showerhead pressure ration or blowing ratio of the prior art. Another object of the present invention is to improve the efficiency of a gas turbine engine by providing improved cooling for the leading edge region of the airfoil.